This invention relates generally to gas turbine engine turbines and more particularly to apparatus for sealing turbine sections of such engines.
A gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. In a turbojet or turbofan engine, the core exhaust gas is directed through an exhaust nozzle to generate thrust.
A turbofan engine uses a low pressure turbine downstream of the core to extract energy from the primary flow to drive a fan which generates propulsive thrust. The low pressure turbine includes annular arrays of stationary vanes or nozzles that direct the gases exiting the combustor into rotating blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. Typically two or more stages are used in serial flow relationship.
These components operate in a high temperature environment. Nearby components outside the gas flow path (such as casings) must be protected from the high temperatures to ensure adequate service life. Leakage of flowpath gases between components, for example between turbine rotor shrouds and adjacent turbine nozzles, is therefore undesirable. Prior art designs have attempted to minimize the leakage gap through the compression of the honeycomb on the shroud. While somewhat effective this does not completely prevent leakage.
Accordingly, there is a need for a turbine shroud configuration that prevents leakage between the shroud and adjacent components.